High pressure turbine airfoil recovery device and method of heat treatment

ABSTRACT

A fixture and method for repairing superalloy articles. The fixture ideally holds the article in place if repairs are made. An interface between the fixture and the article facilitate transfer of heat between the article and the fixture so that the article can be differentially heat treated. A portion of the article extends from the fixture. This portion of the article, which may be repaired within the fixture or may be repaired elsewhere, is solution heat treated while in the fixture so that the area extending from the fixture is solutioned, while heat is transferred from the article through the fixture, thereby preventing the temperature of the portion of the article within the fixture from being elevated so as to modify its microstructure. The solutioned portion can then be heat treated, while in the fixture, to precipitation harden it as desired.

FIELD OF THE INVENTION

The present invention is directed to an apparatus for fixturing anairfoil and to a method for solutioning a portion of the airfoilfollowed by aging the solutioned airfoil

BACKGROUND OF THE INVENTION

An aircraft gas turbine engine or jet engine draws in and compresses airwith an axial flow compressor, mixes the compressed air with fuel, burnsthe mixture, and expels the combustion product through an axial flowturbine section that powers the compressor. The turbine section of theengine includes one or more disks, each disk including a plurality ofblades projecting from its periphery. The hot exhaust gases strike theblades causing the disk(s) to rotate. The rotating disk(s) are attachedto a shaft that also drives a compressor. The compressor is also madefrom rotating disks, each disk having a plurality of blades projectingfrom its periphery. The disk turns rapidly on a shaft as the shaft isrotated by the turbine, and the curved blades draw in and compress airin somewhat the same manner as an electric fan.

The turbine blades on the disk are in the hot exhaust gases resultingfrom the combustion of the fuel and rotate at very high speeds. Thus theblades operate in an oxidative and corrosive environment, and aresubjected to high operating stresses. In order to survive these harshconditions, the turbine blades are made from superalloys, an expensiveblend of elements that provide oxidation resistance, corrosionresistance and strength. These superalloys are further strengthened inpreferred directions by various mechanisms which include growing theturbine blades as directional grains or even as single crystals.

The superalloys used for turbine blades include nickel-basedsuperalloys, iron-based superalloys and cobalt-based superalloys. Thesesuperalloys can be further strengthened by precipitation mechanisms. Forexample, gamma prime (γ′) phases comprising Ni₃Al are precipitated inthe gamma matrix of the FCC crystal structure of the alloy byappropriate solutioning and aging treatments. Controlling the gammaprime phases, both the size and distribution for these as-cast, newparts is well-known. The turbine blade comprises an airfoil portion thatextends into a hot gas stream, a dovetail portion that attaches theblade to the turbine disk and optionally a platform portion thatseparates the airfoil portion from the dovetail portion. A shank portionis intermediate the platform portion and the dovetail portion. Theturbine blades are also provided with environmental coatings and/orthermal barrier coatings to further improve their survivability in thehot, corrosive, oxidative environment of a turbine engine.

The turbine blades nevertheless are subject to damage as a result ofoperation in the gas turbine engine. This damage can be both mechanicalin nature as well as metallurgical in nature. The turbine blades areexpensive to produce, so that it is desirable from an economicstandpoint to repair the blades rather than replace them wheneverpossible. In many situations, the blades can be repaired by removing anyremaining protective coatings, followed by welding damaged mechanicalareas and reworking the weld repaired areas to restore the dimensions asrequired, followed by reapplication of the protective coatings.

The metallurgical damage is inherent as a result of normal operation ofthe gas turbine engine. The blade material, typically a high temperaturesuperalloy, derives its corrosion and oxidation resistance by selectionof a careful combination of elements. Its strength is a result offormation of fine precipitates by precipitation hardening. However,operation of the blades at the high temperatures of the turbine enginefor extended periods of time results in fatigue, creep-rupture andrafting. Rafting, when it is present, occurs in the turbine blade fromthe platform to the top, and more specifically, in the blade trailingedge below the tip and above the platform. As used herein, rafting meanscoarsening of the γ′ precipitate and precipitate alignment in thedirection of applied stress. This region having the coarsened γ′ ischaracterized by reduced strength.

Welding of the blades to repair mechanical damage is known. The damagerepaired by welding is not limited to blades removed from engineservice, but can also be new blades requiring repair of casting defectsor damage resulting from testing or machining. Although various weldrepair methods are available, one illustrative weld repair method isSWET welding, that is, Superalloy Welding at Elevated Temperatures. Thisprocess was developed by General Electric Aircraft Engines in the 1980'sto commercially repair damaged turbine blades refurbished for itsairline customers. The process entails heating of the blade airfoil toan elevated temperature using a heat source such as quartz lamps orinduction coils. The heat source is focused in a narrow region of theblade undergoing repair, typically at the blade tip, and heat islocalized at the blade tip. An area of the blade to be repaired is thenwelded while maintaining the blade at elevated temperature. A variationof SWET welding techniques appears in various prior art references.

The problem with welding of blades is that the weld area and heataffected zone (the “HAZ”) are heated above the solutioning temperatureof the gamma prime. As the blade cools, heat is conducted away from thenarrow area of weld repair and the HAZ, but γ′ will precipitate in thisarea. As heat is transferred away from the repair area, and in portionsof the HAZ, the temperature of the metal is increased, but not to atemperature sufficient to raise the alloy above the solutioningtemperature of the γ′. While fine γ′ precipitate forms in the narrowregion of weld repair and a portion of the HAZ, the γ′ precipitate innarrow adjoining regions further coarsens and is characterized byfurther reduced strength. Thus, mechanical repair by welding does notprovide a solution to the metallurgical problems related to extended useof precipitation-hardened alloys at elevated temperatures, and incertain cases, may further exacerbate the problem.

Although the prior art discloses that the weld area is desirably stressrelieved and rapidly cooled to a temperature below the γ′ precipitationhardening temperature, the prior art does not address the problem ofrafting in other regions of the airfoil portion of the blade that may bedistal weld repair. The prior art references indicate that, in order toreduce the residual stresses in the repaired article, the weld-repairedblade (i.e. the entire blade) is placed in a fixture and stress relievedin a furnace, as is typically done on weld-repaired articles subject tostress, although the written description is otherwise devoid of adiscussion of how post-weld treatments are accomplished. The prior artdoes not recognize the need to restore the metallurgical properties ofairfoils not subject to weld repair.

A method for sintering a wear resistant layer to a blade tip isdisclosed in U.S. Pat. No. 4,818,833 issued Apr. 4, 1989. This patentidentifies utilizing a radiant heat source to heat the blade tip. Theradiant heat source is a graphite susceptor having heating chamberswhich extend into the susceptor. No post-sintering heat treatments arediscussed; however the patent does disclose coarsening of the gammaprime near the blade tip that was judged to be acceptable, but does notrecognize the problem of continued coarsening of this gamma prime thatwill inevitably result from high temperature exposures for long periodsof time. Of course, obvious methods of heating may be substituted forthe radiant heat source, such as for example, induction coils, but suchsubstitutions do not provide recognition of the problem of continuedcoarsening, or alleviation of coarsening that has already occurred.

U.S. Pat. Nos. 6,020,571 (the '571 patent) and 6,124,568 (the '568patent), assigned to the assignee of the present invention andincorporated herein by reference in their entirety disclose acceptablemethods of welding of precipitation hardenable nickel-based superalloyblades by carefully controlling heat input during the welding process.While effective for advancing the art of welding, the '568 patent onlydeals with the problem of γ′ precipitation resulting from the weldingoperation, and this problem is controlled by controlling the heat inputin the localized region undergoing repair in accordance with a desiredtemperature profile.

In order to achieve a uniform γ′ precipitate structure throughout theblade, either as a result of post-welding coarsening or as a result ofrafting due to extended exposure to high temperatures, one currentpractice is to solution the entire blade, thereby dissolving all γ′,followed by reprecipitation of the γ′. This process is effective inproviding uniform γ′ in the airfoil portion of the blade, but presentsother problems. Specifically, dovetail regions of the blade and theshank regions of the blade, both located below the platform region, aresubjected to high stresses, either from machining operations or fromservice induced stresses as a result of being positioned in the disk athigh rotational speeds, or both. These operations result in residualstresses. The portions of the blade below the platform also experiencean operating temperature significantly lower than the airfoil portion ofthe blade, which extends into the hot gas stream. Thus. a hightemperature solutioning treatment of those portions of the blade belowthe platform is undesirable as it can result in recrystallization ofthese portions due to the residual stresses. Since many modern bladesare either directionally solidified (providing large columnar grainsoriented parallel to the longitudinal axis of the blade) or aresolidified as single crystals, recrystallization in this region isundesirable as it reduces strength of the blade in this region, whichcan be a limiting factor for the mechanical properties of the entireblade.

What is needed is a technique that permits restoration of metallurgicalproperties of the airfoil portion of a blade having γ′ precipitates toeliminate the problem of rafting. The rafting can be a result ofextended use of the blade at elevated temperatures and load. Therestoration of the metallurgical properties of the airfoil portion ofthe blade should be accomplished without adversely affecting the as-castgrain structure of the portions of the blade at or below the bladeplatform, such as the dovetail area of the blade.

SUMMARY OF THE INVENTION

The present invention provides apparatus and a method for heat treatinga precipitation-hardened article having a thick section and a thinsection so that the thin section can be solution annealed while themetallurgical structure of the thick section is substantially unaffectedby the solution-annealing process. In particular, the present inventionprovides a method for restoring the microstructure of the thin sectionuniformly by solution annealing the thin section to achieve apreselected microstructure but without affecting the microstructure ofthe thick section. More specifically the article is superalloy turbineblade used in a gas turbine engine, the thin section is the airfoilportion, which is solution annealed and then heat treated to obtain thedesired microstructure, while the thick section is the portion below theplatform which is maintained at a temperature below itsrecrystallization temperature.

A turbine blade used in a gas turbine engine comprises an airfoilportion extending outward into the hot gas flow path of the turbineengine, a dovetail portion that attaches the turbine blade to a turbinedisk and a platform portion that is intermediate the airfoil portion andthe dovetail portion. A shank portion provides a transition between theplatform portion and the dovetail portion, the shank portion and thedovetail portion comprising the blade below the platform portion. Theairfoil portion of the turbine blade is of a thin, curvilinear designengineered to allow for smooth fluid flow of gas over the airfoilportion surface in the engine gas flow path, while maximizing the energywithdrawn from the hot gas. The gas contacting an flowing over theairfoil portion powers the turbine. A plurality of turbine blades areattached to the turbine disk. While the platform is an optional featurein a blade, the platform is oriented substantially perpendicular to thedovetail portion and the airfoil portion. The platform assists inreducing the leakage of hot corrosive, oxidative combustion gases belowthe platform and between the blade dovetail portion and the mating diskdovetail slots, thereby providing protection to both the disk and theblade dovetail portion from the hot corrosive/oxidative combustiongases.

The present invention accomplishes restoration of the metallurgicalstructure of the airfoil portion of the turbine blade to that whichapproximates the microstructure of a new-make blade, typically anas-cast microstructure, without disturbing the microstructure of theblade below the blade platform, thereby restoring the fatigue andcreep-rupture properties to the blade, as well as other mechanicalproperties. The article and method of the present invention permitsrestoration of blades removed from service that have deterioratedmechanical properties resulting from rafting, as well as blades thathave been weld repaired to restore damaged portions of the airfoilportion. The weld repair may be accomplished by any traditional weldrepair method, such as SWET welding, TIG welding laser welding or anyother well-known welding repair method to restore the airfoil portion ofdamaged new-make blades or blades removed from service. The bladeundergoing metallurgical restoration, as discussed above, experiencesrafting in the airfoil portion either as a result of in-serviceoperation or as a result of weld repair, or both. In order to restorethe microstructure of the blade to new-make conditions, typically anas-cast microstructure, rafting is eliminated from the airfoil portionby heating only the airfoil portion above the solutioning temperature,thereby resolutionizing it. The airfoil portion can then beprecipitation-hardened to provide a preselected microstructurecompatible with its intended future use.

The microstructure restoration of the present invention discussed aboveis accomplished by placing the turbine blade in the fixture of thepresent invention. The fixture comprises a high conductivity material.The fixture includes a receptacle having a geometry for receiving theportion of the blade below the blade platform, at least the dovetailportion, and typically the shank as well. The portion of the blade belowthe platform is placed in the receptacle of the fixture so that heatfrom the airfoil portion solutioning operation can be transferred awayfrom the blade below the platform by operation of the fixture. A portionof the blade, typically including the airfoil portion of the blade,extends from the fixture. The fixture further includes a means forremoving heat from the blade and fixture while preventing fixture fromoverheating. Typically, a cooling fluid contacts a surface of thefixture to transfer heat away from the portion of the blade below theplatform. Because the solutioning temperature for each specific alloywill vary, the solutioning temperature is dependent on alloy compositionas well as the size and distribution of the γ′ desired for a givenapplication, but the solutioning temperature and time are selected so asto solution γ′ precipitates. To further elaborate on the variability ofsolutioning temperature, Rene′ N5 is a well-known superalloy compositionused for turbine blade applications having a range of compositions. Thesolutioning temperature for alloy compositions falling within the Rene′N5 range, which has a compositional range in weight percent of 5-10percent cobalt, 5-10 percent chromium, 5-7 percent aluminum, 3-8 percenttantalum, 3-8 percent tungsten, up to 2 percent molybdenum, up to 6percent rhenium, 0.08-0.2 percent hafnium, 0.03-0.07 percent carbon,0.003-0.006 percent boron, up to about 0.02 percent yttrium and thebalance nickel, can vary significantly. The solutioning temperature willvary as the compositional range of the alloy changes, even within theidentified range limits. This is true as well for other superalloystypically manufactured within an identified compositional range. Formost nickel based superalloy blades, the solutioning temperature is inthe range of about 1900-2400° F. for a time of about 0.25-24 hours.

While maintaining a supply of cooling fluid across a surface of thefixture, the portion of the blade extending from the fixture is heatedunder a protective atmosphere. The blade can be heated by selectiveapplication of heat to a portion of the blade extending form thefixture, typically its airfoil portion, using any suitable means forheating. The airfoil portion of the blade is heated to a temperature inthe solutioning temperature range of the alloy. The fixture conductsheat from the platform and the portion of the blade below the platform,so that the temperature rise is sufficiently low in the portion of theblade below the platform that neither recrystallization nor coarseningof the precipitates occurs. This temperature of the airfoil portion ismaintained for a time sufficient to solution intermetallic phasespresent in the form of precipitates. Then, the heat source is removed.The fixture quickly transfers heat from the airfoil portion of theblade, precipitates forming in the airfoil portion as it cools. If theprecipitate size is not proper, heat can be applied to the airfoilportion of the blade to permit coarsening to the proper size, asrequired.

The solutioning temperature is an inherent characteristic temperature ofeach precipitation-hardenable alloy. This solutioning temperature isreadily available for all commonly used precipitation-hardenable alloys,and is readily determinable for new precipitation-hardenable alloys anduncatalogued precipitation-hardenable alloys by one skilled in the art.The solutioning temperature is an elevated temperature at whichprecipitates, typically a separate intermetallic phase formed at anaging temperature below the solutioning temperature, dissolve in themetal matrix yielding a uniform matrix substantially free of γ′.Material systems having more than one intermetallic phase may have morethan one solutioning temperature, as intermetallic phases of differentcompositions dissolve at different elevated temperatures. Solutioningtemperatures of nickel base superalloys are in the range 1900-2400° F.Precipitation-hardening temperatures are below the solutioningtemperature and can be as low as 1400° F., but will vary depending onthe alloy and the size of the precipitate. The rate of formation andgrowth of precipitates in a matrix is a function of time at temperatureas is known in the art.

Other features and advantages of the present invention will be apparentfrom the following more detailed description of the preferredembodiment, taken in conjunction with the accompanying drawings whichillustrate, by way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 depicts a perspective view of a repaired airfoil.

FIG. 2 depicts a front view of a repaired airfoil assembled to thefixture of the present invention.

FIG. 3 depicts a front view of the fixture of the present invention usedto solution heat treat airfoils in accordance with the method of thepresent invention.

FIG. 4 depicts a perspective of an airfoil in a fixture with a heatsource applied to the airfoil in accordance with the best mode ofpracticing the method of the present invention.

FIG. 5 depicts a front view of a repaired airfoil assembled to a secondembodiment of the fixture of the present invention having capabilitiesof cooling the internal cavity of the airfoil.

DETAILED DESCRIPTION OF THE INVENTION

The present invention provides a method for solutioning the airfoilportion of a turbine blade and then uniformly age hardening the airfoilportion of the turbine blade without recrystallizing the portion of theturbine blade below the platform. This restoration is required once themicrostructure of the airfoil portion has been modified. Thissolutioning and age hardening treatment of the airfoil portion of theturbine blade restores mechanical properties uniformly to the airfoilportion of the blade, while restoring a substantially uniformmicrostructure. The present invention further provides apparatus in theform of a fixture for accomplishing the solutioning and subsequent agehardening of the airfoil portion of the turbine blade.

A typical turbine blade 2 is depicted in FIG. 1. However, the inventionis not limited to a turbine blade, and may include other configurationshaving a thin section and a thick section wherein the microstructure ofone section must be modified without significantly altering themicrostructure of the other section. Another such structure within aturbine engine is a turbine vane also generically defined as a turbineairfoil, which also has a thin section extending from a thicker section.This structure is distinguished from the turbine blade in that it issubstantially stationary, being a non-rotational part. However, the thinsection still may require microstructure restoration. Turbine blade 2includes an airfoil portion 4 extending from a platform portion 8. Theportion of the blade extending from the platform 8 in a directionopposite the airfoil portion includes a dovetail portion 6 and a shankportion 9 intermediate the platform portion 8 and the dovetail portion6. The airfoil portion 4 further includes a tip section 10, a leadingedge 12 and a trailing edge 14. In FIG. 1, the tip region 10 has beenweld repaired to restore missing metal thereby restoring the dimensionsof the blade. Tip region 10 is the region most frequently repaired.However, the weld repair may occur anywhere along the airfoil portion 4,the platform portion 8 or the seal lips. Welding in the tip region isassociated with rafting in the area adjacent to the weld repair.However, rafting, typically resulting from service in a turbine engine,can occur anywhere on the airfoil portion 4 above the platform 8.

FIG. 2 depicts turbine airfoil 2 inserted in fixture 20. FIG. 3 depictsfixture 20 without a turbine airfoil 2 inserted therein. The portion ofthe blade below the platform, the shank portion 9 and the dovetailportion 6, is inserted into fixture 20, which includes a dovetail slot22 corresponding to blade dovetail portion 6. In the embodiment depictedin FIG. 2, the fixture further includes an orifice 24 extendingtherethrough, the orifice 24 having an inlet 26 and an outlet 28. Asshown, the inlet 26 and outlet 28 are depicted as being on the same face30 of fixture 20, but the invention is not so restricted, as the inlet26 and the outlet 28 may be positioned at any convenient locations onfixture 20. The purpose of the inlet and the outlet are to provide acooling means to remove heat from the fixture, so any means to removeheat from the fixture during solutioning may be utilized. In FIG. 2, acooling fluid enters inlet 26, transits through orifice 24 extendingthrough fixture 20, and exits through outlet 28, absorbing heat byconduction.

The receiving area of the fixture 42 in FIG. 4 is a surface configuredto mate with a corresponding surface of a workpiece, here the portion ofturbine blade 2 below the platform portion 8 that includes dovetailportion 6 and shank portion 9. The fixture may be designed so thatplatform portion 8 rests on the top surface 34 of fixture 20. It isintended that the amount of surface contact between these matingsurfaces be maximized to facilitate the transfer of heat from the hottersurface, here the turbine blade as will be explained, to the coolersurface, the fixture. If desired, an optional conductive material 32,such as a conductive grease, can be inserted to fill any gaps betweenthe mating surfaces, thereby increasing the heat transfer between them,as it is well known that air gaps can provide effective insulation whichcan inhibit conduction.

A cooling fluid passes through orifice 24. Cooling inlet 26 and coolingoutlet are adapted to receive a connection to facilitate fluid flow.This may be accomplished by any convenient mechanical connection, suchas a threaded connection, a slip fitting, a compression fitting, afriction fitting, an interference fitting etc. Alternatively, theconnection may be somewhat more permanent, such as a soldered connectionto a line from the fluid source. While the connection should provide afluid-tight connection to prevent leakage of fluid at the connection,the connection is not a primary aspect of the invention.

Fixture 20 is comprised of a highly conductive material that, whenproperly cooled, has a temperature capability consistent with thesolutioning temperature of the article, here turbine blade 2. As turbineblades typically are nickel-based superalloys with solutioningtemperatures in the range of about 1900-2400° F., fixture 20 ispreferably copper. While cooling fluids that can be supplied to orifice24 may include gases or liquids, water is the preferred cooling fluid.However gases, such as nitrogen, or inert gases or other suitableliquids including water solutions may be substituted for water.

Once the repair of the article is accomplished, here the airfoil portion4 of a turbine blade, the blade is placed in fixture 20 so thatconductive contact is established between the article and the fixture.Referring again to FIG. 4, this typically is accomplished by fabricatingthe receiving area 42 of fixture 20 to mate or interface with theadjacent surface 44 of the article, here the portion of the blade belowthe platform that includes the shank portion 9 and blade dovetailportion 6, so that the article 2 is supported when inserted in thefixture 20, while leaving the area requiring solutioned, here airfoilportion 4, exposed. Importantly, the adjacent surface 44 of the article,here the portion of the blade below platform portion 8 that includes theshank portion 9 and dovetail portion 6, and the receiving area 42 of thefixture are in intimate contact to facilitate transfer of heat acrossthe interface. If required to improve conductivity between the articleand fixture 20, a conductive grease 32 can be applied to fill any voidsbetween the receiving area 42 of fixture 20 and mating surface 44 of thearticle, thereby improving heat transfer between the article and thefixture.

As will be recognized by those skilled in the art, fixture 20 providessubstantial support for an article such as a turbine blade 2, so thatany weld repairs to airfoil portion 4 can be accomplished with theturbine blade 2 inserted into fixture 20 if desired.

Once the article, turbine blade 2 is assembled into fixture 20, fluidconduits 43 are established at fixture inlet 26 and outlet 28, andcooling fluid flow is established from a fluid source 45 through fluidconnections 43 and through orifice 24, a protective atmosphere isprovided around the area being heated. The protective atmosphere can bea reducing gas such as nitrogen, an inert gas such as argon, or even apurged vacuum if at least the area to be solutioned is placed into aprotective atmosphere that is purged by application of the vacuum. InFIG. 4, the assembly is placed in an enclosure 61, which is thenprovided with a protective atmosphere. The enclosure may take any form61. However, no enclosure is required, and the protective atmosphere maybe provided by maintaining a positive pressure of the protective gasover the area being solutioned. In a preferred embodiment, the heat canbe applied selectively to the area requiring solutioning, here airfoilportion 4, which is heated to the solutioning temperature of thematerial. As noted previously, the solutioning temperature is aninherent property of the precipitation hardenable material beingsolutioned and is determined by its composition. Turbine blades andvanes typically are nickel-based superalloys having a solutioningtemperature in the range of 1900-2400° F. Heat is applied to the thinarea, the airfoil portion 4, by any convenient means to raise thetemperature of this area to the solutioning temperature, but below themelting temperature of the material. The preferred method of applyingheat to the repaired area is by use of an induction coil 46. FIG. 4depicts three induction coils 46 connected to a power supply 48 via anelectrical cable 50 so as to provide uniform heating of the airfoilportion. In FIG. 4, an appropriate seal may be provided to seal theenclosure where the cable 50 runs through it, but this is not a criticalaspect of the invention, such sealing arrangements being well known tothe art. Heat can be applied to the repaired area by any other means,such as by the radiant energy produced by quartz lamps (i.e. radiantlight) focused on the portion of the blade extending from the fixture,by use of a susceptor, by inserting the article and fixture into afurnace etc. Again, the method of heating is not an important aspect ofthe invention, as methods of uniformly heating a surface whilemaintaining a protective atmosphere are well-known to the art.

As the thin section is solutioned by heating, here the airfoil portion 4of turbine blade 2, heat is transferred through the thick section of thearticle, here the portion of the turbine blade below the platformportion 8, to fixture 20. The flow of cooling fluid, in the preferredembodiment water, through orifice 24 in the copper fixture, aspreferred, transfers heat away from the copper fixture therebypreventing it from overheating. Even though the airfoil portion 4 isevenly heated to an elevated temperature so that it is solutioned, thetransfer of heat through the fixture and to the cooling fluid issufficient to keep the remainder of the blade at a temperaturesufficiently low so that microstructural changes do not occur.Specifically, recrystallization in the worked portions of the bladebelow the platform portion 8 is to be avoided. Thus, even though theairfoil portion is solutioned, the shank portion 9 and the dovetailportion 6 remains relatively cool and the microstructure in theseportions remains substantially unchanged by the processing, whilerafting is eliminated in the airfoil portion 4, which is solutioned.After the solutioning operation is complete, the source of heat 46 isremoved, allowing the solutioned area to cool quickly. The rate at whichthe repaired area is cooled can be controlled by controlling or stoppingthe flow of cooling fluid. Rapid cooling should prevent coarseprecipitates from forming. If the microstructure of the article is notas desired after cooling, the entire article can be heated to apreselected aging temperature for a preselected time to provide thedesired microstructure, provided that the aging temperature is below therecrystallization temperature of the thick portions of the article, herethe shank portion 9 and the dovetail portion 6. This aging temperaturewill form γ′ precipitates of the desired size in the airfoil portion,but the temperature is sufficiently low that the microstructure of thedovetail is not affected. Alternatively, the dovetail portion can becooled in the fixture as described above, while the portion of the bladeextending from the fixture is aged for a preselected period of time at apreselected temperature to achieve the desired precipitate size, asdescribed above. Of course, the aging temperature is maintained wellbelow the solutioning temperature Again, as is recognized by thoseskilled in the art, this temperature/time to achieve a desiredmicrostructure size is an inherent characteristic of the material.

After the thin section of the article has been aged to achieve thedesired microstructure, it is suitable to be placed into service. Anyrafting that may have been present in the thin section, here airfoilportion 4, is removed by the solutioning and homogenization of the thinportion. Subsequent development of the precipitates by age hardeningrestores the mechanical properties, and in particular creep-rupture andfatigue properties, to the blade, so that it can be returned to servicehaving mechanical properties equivalent to a new blade.

FIG. 5 depicts an alternate embodiment of the present invention. Becauseturbine blades, particularly high pressure turbine blades, operate atvery high temperatures, both passive and active cooling methods areutilized to prevent overheating. Passive cooling methods include thermalbarrier coating systems. Active cooling is provided by providing theturbine blades with internal cavities or internal cooling passages.These cooling passages are also referred to as serpentine passages andare well known in the art. Cooling air from the compressor is passedthrough these serpentine passages to help to maintain the temperature ofthe blade within temperature limitations. In order to protect the wallsof these internal cooling passages from corrosion and oxidation at theelevated temperatures of operation, these walls are provided with aprotective coating. While the protective coating may be any protectivecoating compatible with the airfoil alloy, one well-suited fornickel-based superalloy turbine blades is an aluminide coating, which isapplied by circulating a vapor phase aluminide through the passage ways,the aluminide being deposited on the walls of the passageways. Thecooling air circulated through the serpentine passageways of the bladealso avoids incipient melting of the thin aluminide coating applied overthe nickel-based superalloy substrate of the internal passageways.

One of the problems faced by blades with internal serpentine passagewayshaving walls protected by aluminide coatings is that the hightemperatures of solutioning can cause the incipient melting of the thinaluminide coating, thereby destroying the protective coating on theserpentine walls and making the internal walls susceptible to oxidationand perhaps corrosion at elevated operating temperatures. However, asshown in FIG. 5, which is a modification of the fixture disclosed inFIGS. 2, 3 and 4, this problem is solved by utilizing the serpentinepassageways during solutioning. FIG. 5 depicts a blade having at leastone internal passageway 70. As shown, the at least one internalpassageway 70 extends through the portion of the blade below platformportion 8, that is the dovetail portion 6 and the shank portion 9. Theat least one internal passageway 70 connects to serpentine passageway 80in airfoil portion 4. In this embodiment, a cooling gas is provided andflows through a gas attachment 72 positioned in fixture 20. A coolinggas, typically an inert gas such as argon is provided to gas attachment72. Gas attachment is in fluid communication with gas vent 74 which, inturn is in fluid communication with the at least one internal passageway70. Appropriate sealing arrangements can be provided between bladedovetail portion 6 and gas vent 74 so that a positive flow of gas ischanneled into the at least one internal passageway 70. Alternatively,fixture 20 can be designed with a movable closure along its verticalwalls so that the blade below the platform portion 8 is sealed infixture 20 after installation, and gas supplied to the fixture 20through gas attachment 72 and gas vent 74 or through a separate gassupply means enters into the at least one internal passage 70. The gasflowing through the at least one internal passage enters into theserpentine passageways 80 of airfoil portion 4 and exits the airfoilportion 4 through apertures 82 located along the trailing edge 14.

The flow of cooling gas through the at least one internal passage 70,into serpentine passageways 80 and out the apertures 82 located alongthe trailing edge 14 provide cooling to the both the at least oneinternal passageway 70 and serpentine passageways 80 during thesolutioning operation, thereby keeping the aluminide coating below itsincipient melting temperature. The incipient melting temperature of anyaluminide coating/nickel-based superalloy combination will varydepending upon the composition of the nickel-based superalloy. Incipientmelting of an aluminide coating applied to a nickel-based superalloy cangenerally be avoided by maintaining the temperature within the coolingpassages below about 2100° F. However, it will be recognized by thoseskilled in the art that in certain nickel-based superalloy/aluminidecoating combinations, incipient melting is not a problem until highertemperatures. For example, aluminide coatings in Rene 142 and Rene N5,two well-known turbine superalloys, incipient melting will be avoided ifthe temperature within the cooling passageways is maintained below about2200° F. Thus, restoration of the microstructure of turbine blades withinternal cooling cooling passageways or serpentine channels can beaccomplished without damaging the aluminide coating applied to the bladeinternal walls by avoiding incipient melting of the applied coatingusing a slight modification of the fixture of the present invention.

While the invention has been described with reference to a preferredembodiment, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment disclosed as the best modecontemplated for carrying out this invention, but that the inventionwill include all embodiments falling within the scope of the appendedclaims.

1. Apparatus for restoring the microstructure of a portion of an articlehaving a thin section and a thick section, comprising: a heat source,the heat source to heat at least a portion of the article to apreselected temperature; a source of cooling fluid; a fixture forsupporting the article, the fixture further comprising a material ofhigh conductivity, a receptacle corresponding to a surface of thearticle for receiving a portion of the article and providing a means fortransferring heat between the article, when inserted, and the fixture,an orifice having an inlet and an outlet, the inlet connected to asource of cooling fluid which removes heat from the fixture; and aprotective atmosphere applied around at least the portion of the articleheated to the preselected temperature, wherein cooling fluid passingthrough the orifice removes heat from the fixture, the heat having beentransferred from the article to the fixture.
 2. The apparatus of claim 1wherein the fixture comprising a material of high conductivity includesa material selected from the group consisting of copper and copper-basedalloys.
 3. Apparatus for solutioning a superalloy turbine blade, theturbine blade having an airfoil portion, a platform portion, a shankportion below the platform portion and a dovetail portion below theshank portion, comprising: means for heating the airfoil portion of theturbine blade to a first preselected temperature in the solutioningtemperature range of the superalloy blade; a source of cooling fluid; afixture for supporting the turbine blade, the fixture further comprisinga material of high conductivity, a receptacle having a geometrycorresponding to a geometry of the turbine blade below the platformportion for providing surface contact with the turbine blade below theplatform portion, when inserted therein, and wherein the airfoil portionof the turbine blade extends away from the fixture when the turbineblade is inserted into the receptacle, an orifice having an inlet and anoutlet, the inlet connected to the source of cooling fluid which removesheat from the fixture; and a protective atmosphere surrounding at leastthe airfoil portion of the turbine blade when the airfoil portion isheated to the preselected temperature, wherein the portion of the bladeinserted in the receptacle is maintained at a second temperature belowthe first preselected temperature by the cooling fluid passing throughthe orifice of the fixture.
 4. The apparatus of claim 3 wherein thefixture comprising a material of high conductivity includes a materialselected from the group consisting of copper and copper-based alloys. 5.The apparatus of claim 3 wherein the means for heating to a firstpreselected temperature includes a heating means selected from the groupconsisting of radiant light, an induction coil and a furnace.
 6. Theapparatus of claim 3 wherein the source of cooling fluid is water. 7.The apparatus of claim 3 wherein the protective atmosphere is anon-reactive atmosphere.
 8. The apparatus of claim 3 wherein the turbineblade further includes internal cooling passages, and the apparatusfurther includes means for providing cooling gas to the internal coolingpassages.
 9. A weld repaired differentially solution treated and agedsuperalloy turbine blade, comprising: a microstructure having apreferred crystallographic orientation; an airfoil portion, a platformportion, a shank portion below the platform portion and a dovetailportion below the shank portion; an area of weld repair in the airfoilportion; the portion of the blade below the platform portioncharacterized by an as-cast microstructure having a substantial absenceof recrystallization; and wherein the airfoil portion has asubstantially preselected uniform size and distribution of γ′precipitates further characterized by a substantial absence of raftingand wherein the area of weld repair and a heat affected zone adjacent tothe area of weld repair are characterized by a substantial absence ofrafting.
 10. The weld repaired superalloy turbine blade of claim 9wherein the preferred microstructure of the blade is selected from thegroup consisting of a single crystal grain and directionally solidifiedcolumnar grains.
 11. The weld repaired superalloy turbine blade of claim9 wherein the area of weld repair is located in the airfoil portion. 12.The weld repaired superalloy turbine blade of claim 11 wherein the areaof weld repair is located in a tip region of the airfoil portion.
 13. Adifferentially solution treated superalloy turbine blade, comprising: amicrostructure having a preferred crystallographic orientation; anairfoil portion, a platform portion, a shank portion below the platformportion and a dovetail portion below the shank portion; the portion ofthe turbine blade below the platform portion characterized by asubstantially as-cast microstructure having a substantial absence ofrecrystallization; and wherein the airfoil portion of the differentiallysolution treated turbine blade has a substantial absence of rafting. 14.The differentially heat-treated turbine blade of claim 13 furtherincluding internal cooling passageways coated with an aluminide coating,the aluminide coating substantially free of incipient melting.
 15. Amethod for using a fixture to differentially heat treat a superalloyturbine blade, comprising the steps of: providing a turbine blade havingan airfoil portion, a platform portion, a shank portion below theplatform portion and a dovetail portion below the shank portion;providing a source of cooling fluid; providing a fixture for supportingthe turbine blade, the fixture further comprising a material of highconductivity, a receptacle corresponding to at least a part of a bladesurface below the platform portion of the turbine blade, the receptacleproviding surface contact with at least the part of the blade surfacebelow the platform surface when inserted therein so that at least theairfoil portion of the blade projects away from the fixture, an orificehaving an inlet and an outlet, the inlet connected to the source of thecooling fluid; inserting at least a part of the blade surface below theplatform portion into the receptacle of the fixture; establishing a flowof cooling fluid from the source of cooling fluid through the orifice;providing a heating source; utilizing the heating source to heat theairfoil portion of the blade projecting away from the fixture; andheating the airfoil portion of the blade to a solutioning temperature ofthe superalloy for a time sufficient to solution gamma primeprecipitates, while maintaining at least the part of the blade below theplatform in the receptacle of the fixture below the solutioningtemperature, the fixture conducting heat away from the blade, and thecooling fluid conducting heat away from the fixture.
 16. The method ofclaim 15 further including the additional steps of: after solutioning,heating the blade to an aging temperature below the solutioningtemperature, the aging temperature and time selected to produce auniform distribution of γ′ precipitates of a preselected sizecharacteristic of the alloy in the solutioned portion of the blade,while not substantially affecting the size and distribution of γ′ in theportion of the blade below the platform.
 17. The method of claim 15further including the additional step of providing a protectiveatmosphere over at least the airfoil portion of the blade during thestep of heating to a solutioning temperature.
 18. The method of claim 15wherein the additional step of providing a protective atmosphereincludes providing a protective atmosphere selected from the groupconsisting of an inert gas and a reducing gas.
 19. The method of claim15 further including the step of applying a high temperature conductivegrease between the receptacle and the part of the blade below theplatform portion in the receptacle, thereby increasing conductivitybetween the blade and the fixture.
 20. The method of claim 15 whereinthe step of providing a source of cooling fluid includes providingwater.
 21. The method of claim 15 wherein the step of heating at leastthe airfoil portion of the blade to a solutioning temperature of thesuperalloy for a time sufficient to solution γ′ precipitates includesheating to a temperature in the range of from about 1900°-2400° F. for atime of about 0.25-24 hours.
 22. The method of claim 15 wherein the stepof providing a fixture comprising a material of high conductivityincludes providing a fixture selected from the group consisting ofcopper and its alloys.
 23. The method of claim 15 wherein the steps ofproviding a heating source and utilizing the heating source to heat theportion of the blade projecting away from the fixture includes providingquartz lamps and focusing the quartz lamps on the portion of the bladeprojecting away from the fixture.
 24. The method of claim 15 wherein thesteps of providing a heating source and utilizing the heating source toheat the portion of the blade projecting away from the fixture includesproviding an induction coil and placing the induction coil around theportion of the blade projection away from the fixture.
 25. The method ofclaim 15 wherein the step of providing a turbine blade further includesproviding a turbine blade having internal cooling passageways, the stepof providing a fixture further includes providing a fixture having ameans for providing cooling fluid to the internal cooling passageways ofthe turbine blade, and the step of establishing a flow of cooling fluidfurther includes establishing a sufficient flow of cooling fluid throughthe means for providing cooling fluid to the internal coolingpassageways so that the temperature of the internal passageways ismaintained below an incipient melting temperature of an aluminidecoating applied to the internal passageways of the superalloy turbineblade substrate.
 26. The method of claim 25 wherein the step ofestablishing a flow of cooling fluid further includes providing asufficient flow of cooling fluid to maintain the temperature of theinternal passageways below about 2100° F.
 27. The method of claim 25wherein the step of providing a superalloy turbine blade includesproviding a nickel-based superalloy turbine blade selected from thegroup consisting of Rene 142 and Rene N5, and wherein the step ofestablishing a flow of cooling fluid further includes providing asufficient flow of cooling fluid to maintain the temperature of theinternal passageways below about 2200° F.